Double frangible bearing support

ABSTRACT

A double frangible bearing support structure supports a low pressure rotor of an aircraft engine. The support structure has a first bearing assembly including a first bearing supported by a first bearing support adapted to buckle or frange when subject to a predetermined critical load resulting from an abnormal rotor imbalance. The support structure has a second bearing assembly comprising a second bearing having rolling elements disposed between inner and outer races. The outer race is connected to a second bearing support by means of frangible bolts adapted to fail when subject to a predetermined critical load resulting from radial displacements and loads of the low pressure rotor following decoupling/franging at the first bearing support.

TECHNICAL FIELD

The application relates generally to aircraft engines and, moreparticularly, to a bearing support arrangement for a spinning rotor inan aircraft engine.

BACKGROUND OF THE ART

The fans of aircraft engines are designed to resist damage caused byforeign object ingestion. However, in certain circumstances, a fan maybe damaged to such an extent that parts of one or more of the fan bladesbecome detached from the rotor disk (referred to herein as a fan bladeoff event or FBO event). This may result in a significant imbalancerequiring shutdown of the engine to minimise load transmission to theaircraft. The imbalance in the fan created by the blade loss generatesextremely high radial loads which must at least be partially absorbed asthe engine is run down to windmilling speed (i.e. the speed at which therotor spins in a non-operative condition as a result of the aircraftmoving through the air).

Under certain circumstances, the vibration resulting from the fanimbalance at windmilling speed can still be considerable. If notappropriately controlled, these vibrations may damage the enginestructure and the aircraft and present difficulties to control theaircraft during approach.

SUMMARY

Therefore, in accordance with one general aspect, there is provided aturbofan gas turbine engine comprising a rotor including a shaft coupledto a propulsive fan, a plurality of axially spaced-apart bearingassemblies supporting the shaft on a stator structure of the engine, theplurality of bearing assemblies including a first bearing assemblyadjacent the rotor, a second bearing assembly adjacent to and spacedapart from the first bearing assembly, and at least a third bearingsupporting an aft portion of the shaft, the first bearing assemblyincluding a first bearing supported by a fusible conical supportstructure, the second bearing assembly including a second bearing havingrolling elements disposed between inner and outer races, the inner racebeing affixed to the shaft, the outer race being connected to a secondbearing support by frangible bolts.

In accordance with a second aspect, there is provided a double frangiblebearing support structure for supporting a low pressure rotor of anaircraft engine, the low pressure rotor including a fan and a lowpressure shaft, the support structure comprising: a first bearingassembly having a first bearing supported by a first bearing supportadapted to buckle or frange when subject to a predetermined criticalload resulting from an abnormal rotor imbalance; and a second bearingassembly comprising a second bearing having rolling elements disposedbetween inner and outer races, the inner race being affixed to the lowpressure shaft, the outer race being connected to a second bearingsupport by means of frangible bolts adapted to fail when subject to apredetermined critical load resulting from radial displacements andloads of the low pressure rotor following decoupling at the firstbearing support.

DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying figures, in which:

FIG. 1 is a schematic cross-sectional view of a turbofan gas turbineengine having a double frangible bearing support arrangement;

FIG. 2 is an axial section view of the double frangible bearingarrangement with a thrust bearing bumper;

FIG. 3 a is an axial section view of the #2 bearing and bumperarrangement; and

FIG. 3 b is a further enlarged axial section view of the #2 bearing andbumper arrangement.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

FIG. 1 illustrates a turbofan gas turbine engine 10 of a type preferablyprovided for use in subsonic flight and generally comprising a lowpressure spool assembly, which includes a fan assembly 12, a lowpressure compressor assembly (not shown) and a low pressure turbineassembly 18 connected by a low pressure shaft 22, and a high pressurespool assembly, which includes a high pressure compressor assembly 14and a high pressure turbine assembly 24 connected by a high pressureshaft 20. The engine 10 further comprises a combustor 16 in whichcompressed air from the high pressure compressor 14 is mixed with fueland ignited for generating an annular stream of hot combustion gasesfrom which the low pressure and high pressure turbine sections extractenergy, as known in the art.

The low pressure spool is rotatably supported by a number of axiallyspaced-apart bearings concentrically mounted about the central axis 11of the engine 10. The low pressure shaft 22 is supported at its front orupstream end by first and second bearings 26 and 28 respectivelycommonly referred to as the #1 and #2 bearings and at a rear end thereofby a third bearing 30 which may be the #5 bearing of the engine (the #3and #4 bearings rotatably supporting the high pressure shaft 20). Thebearing arrangement for a particular engine, including but not limitedto the number and type of bearings selected, is typically determined bya number of factors specific to that engine. The bearing arrangementdescribed herein is exemplary only, and not intended to be limiting. Inthis example, the forward and rearward most bearings, i.e. the #1 and #5bearings, may be roller bearings for radially supporting the lowpressure shaft 22. The #2 bearing 28 may be a thrust bearing to provideboth axial and radial support to the low pressure shaft 22. As shown inFIG. 1, the first bearing 26 is disposed adjacent the fan rotor, whilethe #2 bearing 28 is disposed adjacent to and downstream from the firstbearing 26, and upstream of the high pressure compressor 14 relative toa flow direction of the gases through the engine 10. The bearings andsupporting structure flanges may be arranged in any suitable fashion inthe context of the present application.

The first bearing 26 is supported on the stator structure of the engineby a bearing support 27. To minimize the effect of potentially damagingabnormal imbalance loads (e.g. such as caused by fan blade-off-inducedimbalance loads), the bearing support 27 is designed to frange whensubject to a predetermined critical load. As shown in FIG. 2, the #1bearing support 27 may be provided in the form of a fusible conicalstructure connected to an outer race of the #1 bearing. The fusibleconical structure is adapted to fail when subject to a critical loadingcondition to thereby decouple the #1 bearing 26 from the statorstructure of the engine 10. The fusible conical structure decouples the#1 bearing from the static structure of the engine by buckling,collapsing, crimping, yielding or fracturing. The net result is that theradial support provided to the #1 bearing is either eliminated orreduced to a negligible value.

Referring to FIG. 2, it can be seen that the #2 bearing 28 may beprovided in the form of a thrust bearing having rolling elements 30(e.g. ball bearing elements) retained by a bearing cage 32 and disposedbetween inner and outer races 34, 36. The inner race 34 is securelymounted to the low pressure shaft 22. According to the illustratedembodiment, the inner race 34 consists of two half-races 34 a, 34 b(FIGS. 3 a, 3 b). However, it is understood that the inner race 34 couldbe provided as a one-piece component or other suitable configuration.The outer race 36 has a mounting flange 38 projecting radially outwardlyfrom the front end thereof for connection to the #2 bearing support 40forming part of the stator structure of the engine case. The bearingsupport 40 has a Y-shaped cross-section including first and secondaxially opposed frusto-conical portions 40 a, 40 b flaring away fromeach other and an annular mounting flange 40 c projecting radiallyinwardly from the junction of the first and second frusto-conicalportions 40 a, 40 b. The mounting flange 40 c is provided at its innerdiameter with an axially rearwardly projecting spigot 40 d (FIGS. 3 a, 3b) over which the front end portion of the outer race 36 of bearing 28is adapted to be axially slid when loaded in position from the rear endof the engine 10.

A series of circumferentially distributed frangible fasteners, such asfusible or shear bolts 42 or the like, may be used to fasten themounting flange 38 of the outer race 36 to the corresponding mountingflange 40 c of the bearing support 40. In use, the fusible bolts 42 mayshear for structurally decoupling the #2 bearing 28 from the statorstructure of the engine case, and are sized so that shear preferablyoccurs after decoupling at the #1 bearing 26, thereby preventing thetransmission of potentially damaging imbalance forces or other vibratoryforces to the engine case and the airframe in the event of a fan bladeloss or another abnormal fan rotor imbalance event. It is understoodthat the shear bolts 42 are not the only possible type of decoupler.Other suitable types of decoupler could be used as well. For instance, afusable flange or frangible support could be used to disconnect thebearing from the engine case. In the illustrated embodiment, the shearbolts 42 have a weakened zone 42 a (FIGS. 3 a, 3 b) to cause the boltsto fracture when subject to bending moment tensile loads, or shearloads, fatigue loads, other loads, or combination thereof, above apre-determined magnitude, thereby releasing/decoupling the outer race 36and, thus, #2 bearing 28 from the static structure of the engine case.

The radially inwardly extending flange 40 c of the bearing support 40and the radially outwardly extending flange 38 of the #2 bearing 28 forman inverted flange arrangement which provides more flexibility duringnormal engine operations than a conventional back-to-back flangearrangement. The inverted flange arrangement is not as stiff whensubject to axial loads. The inverted flange arrangement slightly flex sowhen pulling with thrust, the bolts 42 are not exposed to as much loadsand bending as they would if they were on a normal back-to-back flangearrangement. As a result, it is possible to reduce the bolt count andstill meet the load cycle fatigue limits for the bolts under normalengine operating conditions. In this way, in the event of a FBO, lessbolts need to be ruptured for decoupling the second bearing 28 from thestator structure of the engine 10, which may beneficially result in amore reliable decoupling system. The inverted flange arrangement mayalso provide a weight saving in some configurations.

Referring more particularly to FIGS. 3 a, 3 b, it can be appreciatedthat a bumper 44 encircles the #2 bearing 28 to limit, in use, theamplitude of radial excursions, and in this example also the distance ofaxial travel, of the bearing (and therefore also the low pressure shaft22), after decoupling of the #2 bearing outer race as described above.The bumper 44 is preferably configured to withstand the post FBO loadsand transmitted to it by #2 bearing so to thereby constrain the radialand axial excursions as described above. The bumper 44 is constructed toresist substantial impact loads both axially and radially during theinitial phase of an FBO event and to then also to survive and containthe axial and radial loads transmitted through the #2 bearing 28 duringwindmilling. According to one embodiment, the bumper 44 is made ofstainless steel.

The bumper 44 has a ring portion 44 a and a mounting flange portion 44 bextending radially outwardly from the outer surface of the ring portion44 a. The bumper 44 is preferably attached to the bearing support 40independently of the second bearing 28 by bolts 46 mounted inregistering holes respectively defined in the radially outer end of theflange portion 44 b of the bumper 44 and in the larger diameter endportion of the second frusto-conical portion 40 b of the bearing support40. The ring portion 44 a of the bumper 44 has a circular radially innersurface disposed in close proximity to a radially outer surface of theouter race 36 of the #2 bearing 28 and defining therewith an annularradial gap 48. The radial gap 48 has a radial size which is sized sothat, in use, the natural vibratory frequency of the low pressure shaft22 as it orbits about the central axis 11 of the engine 10 afterdecoupling at the second bearing 28 is tuned to a desired frequency orto avoid undesired frequency(ies). The radial thickness of the gap 48 ispreferably also sized such as to not excite other modes of the engine oraircraft. It is preferably sized within a range that is large enough todampen the undesired modes and yet not large enough to have the high andlow pressure shafts 20, 22 contact one another or otherwise interactadversely. According to one example, the radial size of the gap 48 isbetween 0.09″ and 0.15″. It is understood that each engine, depending onits mass and resonance (among other factors), would have a differentgap.

Optionally, the axially front end of the ring portion 44 a of the bumper44 has an annular axially forwardly facing surface 44 c which is axiallyspaced from an opposing axially rearwardly facing surface of themounting flange 38 of the #2 bearing 28 by an axial gap 50. The annularaxially facing surface 44 c of the bumper 44 provides an arresting orabutment surface against which, in use, the axially rearwardly facingsurface of the mounting flange 38 may come into contact when movedaxially rearwardly under thrust forces after decoupling. The mountingflange 38 of the outer race 36 is axially trapped between the mountingflange 40 c of the support bearing 40 and the bumper 44. The axial gap50 may be used to limit how far rearwardly the low pressure spool maymove after an FBO event. The axial gap 50 may be sized within a rangethat is large enough to allow the outer bearing race 36 to axially comeoff the spigot 40 d upon rupturing of the frangible bolts 42 and issmall enough to prevent the low spool from gaining too much kineticenergy. Typically, the longer the low pressure spool accelerates afterrupturing of the bolts 42, the more severe the impact with the bumper 44will be. The axial gap 50 is thus preferably kept small to contact thelow pressure spool as soon as possible after decoupling to impedeexcessive acceleration, which may assist the bumper 44 sustain lessloads, and therefore damage, from the initial impact of the secondbearing 28. Also, the axially facing surfaces 38 a, 44 c of the bearing28 and the bumper 44 interact axially to prevent the low pressure shaft22 and the high pressure shaft 20 from contacting each other adverselywhich could eventually result in the rupturing of the low pressure shaft22 and lead to a catastrophic event.

The outer race 36 of the bearing 28 may have a flange 39 projectingradially outwardly from a rear end portion thereof. The rear flange 39is axially spaced from the ring portion 44 a of the bumper 44 by anaxial aft gap 51. As will be seen hereafter, in use the bearing 28 maymove within the axially forward and aft gaps 50 and 51 after decoupling.Axial retention after a FBO event is achieved in this example byentrapment of the bearing outer race 36 between the main bearing support40 and the bumper 44.

The ring 44 a need not be a ring, per se, but rather be any generallyannular assembly suited to functionality described herein, and may besuitably configured as a monolithic structure, a segmented structure, acontiguous annular structure, an interrupted annular structure, and soon. Still other modifications will be apparent to skilled reader.

The reader will also appreciate that axial containment of the LP rotormay be provided by any suitable arrangement and need not be incorporatedinto the bumper assembly as described herein.

From the foregoing, it can be appreciated that the bumper 44 ispositioned relative to the second bearing 28 both limit and control themovement of the bearing outer race, and thus the low pressure shaft 22,after franging.

During normal engine operation, the bumper 44 is inactive. However, inthe unlikely event of an FBO or other exceptional event resulting inexcessive rotor imbalance conditions, considerable radial loads aretransmitted from the low pressure shaft 22 to the first and secondbearings 26 and 28. When these loads reach a critical value, the #1bearing support 27 is configured to buckle or otherwise deform orfrange, thereby decoupling the first bearing 26 from the statorstructure of the engine 10. Since such franging of the #1 bearingsupport 27 has left the upstream end of the low pressure shaft 22 withconsiderably reduced radial support (if any), the upstream end of thelow pressure shaft 22 will tend to start orbiting about its axis 11. Theorbiting of the low pressure shaft 22 and the bending loads on the shaft22 will tend to induce moment loads on the #2 bearing 28. Theseadditional loads on the #2 bearing 28 will tend to rupture the shearbolts 42, releasing the second bearing 28 from the spigot 40 d, therebydecoupling the #2 bearing 28 from the stator structure of the engine 10.Upon decoupling, the second bearing 28 will radially tend to impact uponthe bumper 44, and in this example any axial movement of the rotor willalso be constrained by the described assembly. The axial and radial gaps50 and 48 are, as described above, preferably sized to minimize theinitial impact forces that the bumper 44 has to survive, as well as tuneout unwanted vibratory modes and generally control load transfer to theengine and aircraft. At this point, the #2 bearing 28 is free to orbitin a larger diameter within the rigid ring portion 44 a of the bumper 44preferably such that the resonance of the low pressure shaft as it goesthrough its modes cannot couple to the engine hardware. The radialexcursions and preferably also axial movement of the low pressure shaft22 during windmilling is constrained by the bumper assembly. Theconstraining action of the bumper 44 on the orbiting motion of the lowpressure shaft 22 also preferably impedes intershaft and other harmfulrubbing. The reduced radial support stiffness may be tuned to impede thelow pressure spool from approaching its natural frequency of vibrationto thereby limit transmission of harmful or unwanted loads or vibrationsto the airframe, including cockpit or cabin.

The given examples of the frangible bearing support structure allowreducing the transmission of loads from the LP rotor to the engine andaircraft frame structures in the event of a FBO event. The cabin noiseis addressed through utilization of existing engine hardware, with theexception of the bumper, as opposed to adding specially designedmitigation devices. The buckling/franging of the first bearing supportfollowed by franging the #2 bearing support provides a mechanism bywhich the LP rotor running through the resonance at a lower speed.Avoidance of the selected resonance may be used to limit cockpit, cabinor other airframe accelerations to within desired levels during enginewindmilling.

The above description is meant to be exemplary only, and one skilled inthe art will recognize that changes may be made to the embodimentsdescribed without departing from the scope of the invention disclosed.For example, different materials, different combination of loading inthe fusible structures or bolts, different arrangement of bearings orbearings support flanges, different types of fusible bolts, differentnumbers or sizes of fusible bolts. Still other modifications which fallwithin the scope of the present invention will be apparent to thoseskilled in the art, in light of a review of this disclosure, and suchmodifications are intended to fall within the appended claims.

What is claimed is:
 1. A turbofan gas turbine engine comprising a rotorincluding a shaft coupled to a propulsive fan, a plurality of axiallyspaced-apart bearing assemblies supporting the shaft on a statorstructure of the engine, the plurality of bearing assemblies including afirst bearing assembly adjacent the rotor, a second bearing assemblyadjacent to and spaced apart from the first bearing assembly, and atleast a third bearing assembly supporting an aft portion of the shaft,the first bearing assembly including a first bearing supported by afusible conical support structure, the second bearing assembly includinga second bearing having rolling elements disposed between inner andouter races, the inner race being affixed to the shaft, the outer racebeing connected to a second bearing support by frangible bolts.
 2. Theturbofan defined in claim 1, wherein the fusible conical supportstructure is configured to buckle or frange when subject to apredetermined critical load resulting from an abnormal rotor imbalance.3. The turbofan defined in claim 1, wherein the second bearing is athrust bearing, and wherein after failure of the frangible bolts, thethrust bearing is both axially and radially restrained by a bumper. 4.The turbofan defined in claim 3, wherein the bumper is separatelymounted to the stator structure, the bumper encircling the thrustbearing and having a radially inwardly facing surface disposed in closeproximity to a radially outer surface of the outer race of the thrustbearing and defining therewith a radial gap to accommodate and constrainan orbiting motion of the rotor about the central axis of the engineafter decoupling at the thrust bearing.
 5. The turbofan defined in claim4, wherein the bumper has an axially forwardly facing surface which isaxially spaced by a predetermined axial fore gap from a first flangeprojecting radially outwardly from a front end portion of the outer raceof the thrust bearing, the first flange of the outer race being axiallytrapped between the second bearing support and the bumper; afterdecoupling, the second bearing being free to axially and radially movewithin the radial gap and the axial fore gap.
 6. The turbofan defined inclaim 1, wherein the first bearing has an outer race, and wherein theouter race is supported by the first bearing support.
 7. A doublefrangible bearing support structure for supporting low pressure rotor ofan aircraft engine, the low pressure rotor including a fan and a lowpressure shaft, the support structure comprising: a first bearingassembly having a first bearing supported by a first bearing supportadapted to buckle or frange when subject to a predetermined criticalload resulting from an abnormal rotor imbalance; and a second bearingassembly comprising a second bearing having rolling elements disposedbetween inner and outer races, the inner race being affixed to the lowpressure shaft, the outer race being connected to a second bearingsupport by means of frangible bolts adapted to fail when subject to apredetermined critical load resulting from radial displacements andloads of the low pressure rotor following decoupling at the firstbearing support.
 8. The double frangible bearing support structuredefined in claim 7, wherein the first bearing support comprises afusible conical support structure.
 9. The double frangible bearingsupport structure defined in claim 7, wherein the second bearing is athrust bearing, and wherein after failure of the frangible bolts, thethrust bearing is both axially and radially restrained by a bumper. 10.The double frangible bearing support structure defined in claim 9,wherein the bumper is separately mounted to a stator structure of theengine, the bumper encircling the thrust bearing and having a radiallyinwardly facing surface disposed in close proximity to a radially outersurface of the outer race of the thrust bearing and defining therewith aradial gap to accommodate and constrain an orbiting motion of the rotorabout the central axis of the engine after decoupling at the thrustbearing.
 11. The double frangible bearing support structure defined inclaim 7, wherein the first bearing has an outer race, and wherein theouter race is supported by the first bearing support.